Aircraft lifting rotor and pitch control mechanism therefor



Feb. 23, 1954 HOHENEMSER 2,670,051

AIRCRAFT LIFTING ROTOR AND PITCH CONTROL MECHANISM THEREFOR 3Sheets-Sheet 1 Filed July 18, 1949 INVENTOR.

K4444 HAM m.

Feb. 23, 1954 K H, HOHENEMSER 2,670,051

AIRCRAFT LIFTING ROTOR AND PITCH CONTROL MECHANISM THEREFOR Filed July18, 1949 3 Sheets-Sheet 2 FIG. 6

FIG 8 INVENTOR.

' M A XA Feb. 23, 1954 I K H HOHENEMSER 2,670,051

AIRCRAFT LIFTING ROTOR AND PITCH CONTROL MECHANISM THEREFOR Filed July18, 1949 3 Sheets-Sheet 3 INVENTOR,

KM mww.

Patented Feb. 23, 1954 AIRCRAFT LIFTING ROTOR AND PITCH CGNTROLMECHANISM THEREFOR Kurt H. Hohenemser, Pattonville, Mo.

Application July 18, 1949, Serial No. 105,329

7 Claims.

This invention relates generally to airscrews, and more particularly tolifting rotors for aircraft.

Rotors of this type consist of a plurality of rotary wings or bladesconnected to a rotating hub, and they are arranged to rotate in asubstantially horizontal plane above the fuselage of the aircraft so asto carry its weight. While an airplane requires a certain minimumforward flight speed in order to stay in the air, a rotary wing aircraftis capable of vertical flight.

The present invention is applicable to all types of rotary wing aircraftincluding the helicopter where the lifting rotor provides both lift andpropulsive force, the gyrodyne where separate means of propulsionprovide the propulsive force, the compound rotary-fixed wing aircraftwhere part of the lift in forward flight is provided by a non-rotatinglifting surface and the convertible aircraft which is capable of beingconverted in the air from a rotary wing aircraft into an airplane.

Lifting rotors for aircraft are conventionally driven by engines locatedin the fuselage and connected to the rotor by transmission gears andshafts, or they may be driven by propulsion means located on each of therotary wings, or they are driven like a windmill by the relative airflowacting on the aircraft in forward flight.

In order to carry the weight of the aircraft in vertical flight liftingrotors are required having diameters which are several times greaterthan the diameters of propulsion airscrews for the same size ofaircraft. Because of the tremendous gyroscopic moments acting on suchlarge rigid airscrews during maneuvers of the aircraft it is standardpractice to use hinged lifting rotors where the rotary wings are free toflap vertically up and down.

Hinged lifting rotors however, show a pronounced instability in forwardflight which is increased with increasing flight speed and whichrequires, if not compensated, appreciable skill and constant attentionby the pilot.

A further disadvantage of the hinged lifting rotor is its tendency ofrotary wing stall especially Rotary wing stall causes a reduction andinsevere cases a complete loss of controllability of the aircraft and atthe same time heavy vibrations.

Another disadvantage of the conventional lifting rotor is the failure ofthe rotor to continue rotation and to provide the required lift afterthe driving engines have ceased to operate due to lack of fuel or otherfailure. Hence, in most rotary wing aircraft the pitch angle of therotary wings must be reduced by the pilot by actuating the rotor pitchcontrol in order to secure continued rotation of the lifting rotor inpower-01f flight. At a sufficiently low pitch angle of the rotary wingsrotation is sustained by the'relative airflow acting on the aircraft ina manner similar to the operation of a windmill. If, however, the pilotfails to reduce the rotary wing pitch angle in case of discontinuedengine operation in the air the lifting rotor ceases to turn and toprovide lift, and the aircraft is bound to crash if the rotor is theonly lifting device.-

In view of the generally known unsatisfactory stability characteristicsof the hinged lifting rotor numerous stabilizing devices have previouslybeen proposed. The use of such devices, in most cases, has undesirablesecondary effects, quite apart from their additional Weight and thecomplication of the construction caused thereby. This is true even ofthe simplest of stabilizing devices, the stabilizing tail surface, andin this case the detrimental effects have their origin in the verypowerful turbulent wake of the rotor.

Several proposals have been made for improving the unsatisfactorycharacteristics of the hinged lifting rotor whereby the rotor hub andhinge assembly of a conventional lifting rotor is modified in such amanner as to effect the type of kinematic relation between flappingangle and pitch angle of each rotary wing or between the flapping anglesand pitch angles of certain combinations of several rotary wings of alifting rotor. ever, little was known about the essential parametersaffecting the stability of a rotary wing aircraft, and it may bedemonstrated by the theory of rotary wing flight stability, onlyrecently developed, that neither of the previously proposed improvementsof lifting rotors eliminates the undesirable stability characteristicsof this rotor vp The principal object of the present invention is toprovide an inherently stable lifting rotor so that the aircraft will bestable without the necessity of incorporating additional stabilizingWhen these proposals were made, how? 3 devices with their disadvantageswhereby increasing stability is provided with increasin flight speed.

A further object of the invention is to pro- Vide a lifting rotor whichwill eliminate or reduce rotary wing stall.

Another object of the invention is to provide a liftingrotor which willwindmill after discontinued engine operation in the air withoutrequiring adjustment of the rotor pitch control.

The present invention relates more specifically to the type of liftingrotor where kinematic relations exist between the flapping angles andpitch angles of certain combinations of several rotary Wings of alifting motor. In accordance with the present invention the constructionof the rotor hub and hinge assembly of this type of lifting rotor ismodified in such a manner as to obtain the kinematic relation betweenthe pitch angles and the flapping angles of the rotary wings requiredfor an inherently stable rotor.

The invention will appear more clearly from the following detaileddescription taken in connection with the accompanying drawings, showingbyway of example, preferred embodiments of the invention,

Fig. 1 is a partly schematic side view of a helicopter having an engineinside the fuselage and with a tail rotor provided for the purpose ofcompensating the torque reaction of the main rotor on the fuselage. Theblades are shown in two positions illustrating cyclic flapping, a modeof flapping which will be explained hereafter.

Fig. :2 is a partly schemaic side View of a hellcopter with jet enginesat the tip of the rotary wings. No tail rotor is necessary in this casesince there is no torque reaction of the rotor on the fuselage. Theblades are shown in two positions illustrating collective flapping, amode of flapping which will also be explained hereafter.

Fig. 3 is a plan view of the rotor hub and hinge assembly of a rotortype airscrew where the pitch angle of a rotary wing is increased whenthe wing flaps in the downward direction.

Fig. 4 is a cross section through the wing of Fig. 3 taken on line 44'of Fig. 3. The cross section is shown in two positions in order toillustrate the change of blade pitch angle with the flapping motion.

Fig. 5 is a perspective view of the rotor hub and hinge assembly of alifting rotor showing the preferred embodiment of the invention fora twobladed rotor.

Fig. '6 "is a plan view of the rotor hub and hinge assembly of a liftingrotor showing the preferred embodiment of the invention for. a threebladed rotor.

Fig. '7 is a plan view of the rotor huband hinge assembly of a modifiedrotor embodying the in vention.

.Fig. 8 is a side View .of the walking beam of the .rotor of Fig. 7.

.Fig. 9 is a perspective view, partly schematic, of the rotor hub andhinge assembly of a lifting rotor representing another embodiment of theinvention.

Fig. 10 is a plan view of the .rotor hub and hinge assembly of Fig-.9.

Fig. 11. is a plan view of the rotor hub and hinge assembly of a liftingrotor representing a further-embodiment-of the invention.

Referring now to the drawings in which :like elements are designated bythe same reference characters. throughout all the figures and par-.-

ticularly to Fig. 1, there is illustrated one of the aircraft types towhich the invention may be applied. The helicopter shown is providedwith a single gear driven lifting rotor and with a torque compensatingtail rotor. The rotor shaft I is driven by gears arranged in the gearcase 2, the power being provided by an engine 3. Gear case 2 and engine3 are located inside the fuselage 4 of the helicopter. A transmissionsystem '5 transmits power to the tail rotor 6, which serves tocounteract the torque reaction of the lifting rotor on the fuselage 4.The blades 7 are hinged to the rotor shaft I and are free to flapvertically up and down; The blades 1 are shown in dotted lines in theirmean position. In this posi tion the swept surface of the blades duringrotation forms a slight cone. The resultant rotor force is representedby a dotted vector 8.

The blades 1 are shown in solid lines in a position corresponding to aforward inclination of the rotor. Seen from a point of observationautside the rotor the rotor cone is tilted forward by the angle [3 andthe vector 8 of the resultant rotor force is inclinedforward by the sameangle. Seen from a point of observation whichv rotates with the rotorthe blades flap up and down pe riodically. If the longitudinal axis of ablade points rearward the flapping angle relative to the mean positionis +5 (upward). During rotation of the blade from its rearward positionin a clockwise direction seen from below the flapping angle ,3 isdecreased firstto 0 in the right sidewardposition of the blade and thento '-/3 (downs ward) in the forward position of the blade. In thefurther course of the rotationthe blade;fiap-. ping angle goes againthrough zero the-left sideward position, where the cycle is completed.This kind of flapping motion will becalled cyclic flapping. Seen from anon-rotating point of observation the cyclic flapping motion .of theblades with the amplitude ,5 appears as an in clination of the rotorcone with respect to the mean position by the :anglec.

Fig. v2 illustrates another type of aircraft @to which the invention maybe applied. The hell! ccpter .of .Fig. 2 has a single jet-drivenlifting.rotor. The center part1 of the lifting rotor is: either a, rotatingshaft supported by bearings :inthe fulselage 4,, or it is a non-rotatingstructural member which carries at its upper end-:a bearing to supportthe hub of the lifting rotor.- The blades 1 are again jhingeclzat therotor centerrand: are free to flap up and down. They carry at theirouter ends jet engines 51. Thedotted linesrepresent themeanposition ofthe :blades 1. The solid lines represent "a position of the blades .1

where the flapping angle 18 of all blades is in; creased .by the sameamount. happens for example when the rotational speediof thev rotor:

is reduced while the :rotor'lift is kept constant,

or when the .rotor :lift is increased while the arc-1' tational speed:is kept .constant. A flapping'mm tion where the flapping angles :of-allblades/vary. by the same amount will/be. called collective flapping. Forthe purpose" of this specification the total flapping motion of theblades will be assumed to consist of only two portions: collective:flapping and cyclic flapping. Actually :small flapping oscillationswith higher frequencies. may. occur in addition to these .two .mainflapping modes, but these high frequency oscillations are: insignificantfor the stability .and controlxcharacteristicsrof the aircraft and theyare irrelevant"- for annnderstanding of. the invention.

Figs. 3 and efillustrate the-kinematic relation.

between the flapping angle and the pitch angle of a blade. The directionof rotation of the rotor I is indicated by arrow I0. Links I2 areconnected to the rotor shaft by means of flapping hinges II so as toallow rotation of the links I2 about the axis I3. The blades .1 arerotatably connected to the links I2 so that they may rotate about thelongitudinal blade axis I4 thereby changing the blade pitch angle (Fig.4). This angle is defined as the angle between the direction I ofrotational speed of the blade and the chord axis I6 of the blade sectionas illustrated in Fig. 4.

The blades carry control horns I! having points of attachment I8 forblade pitch control. Moving the points I8 in a vertical directionchanges the blade pitch angle 0. Collective pitch control is achieved bymoving both points I8 simultaneously and by the same amount in thevertical direction. Cyclic pitch control is achieved by oscillating thepoints I3 vertically in opposite direction so that one point I8 goes upwhenthe other point It goes down. It is not necessary foranunderstanding of the present invention to describe the complicatedmechanism required to provide rotor pitch control. In order to avoidconfusion the process of rotor pitch control will be disregarded and itwill be assumed that the points I8 are vertically fixed. This conditionis fulfilled in the neutral position of the cyclic pitch control system.

When the blades 1 are rotated about their longitudinal axis I4 or whenthey fiap about their hinge axis I3, the intersection point IQ of thetwo axes I4 and I3 is fixed in space. Under the assumption outlinedabove the point I8 is the second point of the blade I which isvertically fixed, so that the blade I is only free to rotate about theaxis 20 through the two points I8 and I9. This axis 20 is called thevirtual blade flapping axis. I'he angle between the axes 20 and I3 iscalled 63, and the kinematic relation between the increase in flappingangle A5 and the decrease in blade pitch angle A0 is given by theequation A0=Ae tan a angle 0 is reduced compared with that in the lowerposition (solid lines).

All the lifting rotor arrangements shown in Figs. 1 to 4 are known andthese figures were only included in order to illustrate the meaning ofthe terms which will be used in describing thev invention and to providea better understanding thereof.

Fig. 5 illustrates the hub and hinge assembly of a rotor in accordancewith the preferred embodiment of the invention. A walking beam 2I ishinged in its center to the upper end of the rotor shaft I, therebyallowing a free vertical see-saw motion of the Walking beam about theaxis 22. Links I2 are hinged to the ends of the walking beam 2I soas toallow rotation of the links I2 about the axis 23. Blades 1 are rotatablyconnected to the links I2 by means of a pitch varying pivot so that theymay rotate about their longitudinal axes 14 thereby changing their pitchangle 0. The blades carry control horns I! having points of attachmentI8 for the verrotor so that a description of the control mech-- anism isnot deemed to be necessary.

For a cyclic flapping motion opposite blades flap in oppositedirections. For this type of motion the walking beam 2| rocks about theaxis 22, but no motion takes place about the outer axes 23.

For a collective flapping motion opposite blades flap in the samedirection. For this type of motion the links I2 rock about the axes 23,but

no motion takes place about the center axis 22 of the walking beam 2I.The member2I will also be referred to as a hub like member or as a tippath plane follower since during rotation of the rotor it assumes aposition parallel to the plane defined by the path of the blade tips.

The location of the different axes is more clearly illustrated in Fig.6. The fact that this drawing shows a three bladed rotor does not changethe kinematic relations. Only one of the three blades is illustrated andthe direction of rotation of the rotor is indicated by the arrow I0.Instead of the" walking beam 2I of Fig. 5 a hub 25 with three arms 26 isprovided. The hub 25 is universally hinged to the rotor shaft I asshown. Links I2,

blade I and control arm I! are the same as described before andillustrated in Fig. 5. For the same reason as explained before the hub25 will be also referred to as a tip path plane follower.

For a cyclic flapping motion of the blade I the hub 25 rocks about theaxis 22, but no motion takes place about the outer axes 23. For a col-'lective flapping motion the links I 2 rock about the axes 23, but nomotion takes place about the center of the hub 25. The points I8 areagain' assumed to be fixed in the vertical direction.

In case of cyclic flapping the center point 21 of the axis 22 is thesecond fixed point of the blade I and the blade is only free to moveabout the axis 28 which will be called the cyclic flapping axis andwhich passes through points 21 and I8.

cyclic flapping angle Adm and the corresponding change in blade pitchangle Aacyc is givenby the equation A change in cyclic flapping angleAficyc produces a change in blade pitch angle Aflcyc which isappreciably smaller than the cyclic flapping angle Aficyc.

In case of collective flapping the center point 29 of the axis 23 is thesecond fixed point of the.

blade I and the blade is only free to move about the axis 30 which willbe called the collective.

flapping axis and which passes through points I8 and 29. The anglebetween the axes 23 and 30 is called 63 col and the kinematic relationbetween the increase is collective flapping angle A3501 and thecorresponding decrease in blade pitch angle Aacol is given by theequation The angle between the axes 22 and 2B is called 63 cyaand thekinematic relation between the change in ltnincrease in collectiveflapping angle .Afimi produces a decrease in blade pitch angle Aacolwhich,

quantitatively the correctness of the following ccnclusicns which arebased on qualitative con siderations only, in order to avoid confusion.

A conventional hinged rotor with zero 53 angle has forward level night,when the lifting rotor is power driven, the following main stabilitycharac er st s:

1. increase in forward flight speed produces a Cyclic flapping :of theblades equivalent to abackward inclination of the rotor cone and of theresultant force vector .8 inFig. 1, thereby causing a nose-up momentacting on the aircraft.-

other words the hinged lifting rotor resists an increase in flight speedwhich means, it produces a stabilizing moment when the flight speed ischanged. This is a desirable property of the otor,-

,2, A slow noseeup motion of the aircraft producesa cyclic flapping ofthe blades equivalent to. a backward inclination of the rotor cone andof the resultant force vector 8 in Fig. 1, thereby causing .a furthernoseaup moment acting on the aircraft. In other words, the hingedlifting rotor produces an instabilizing moment, when the atti-- tude ofthe vaircraft is changed. This is a very undesirable property of therotor.

3. A fast nose-up motionof the aircraft with a certain angula speedproduces a cyclic napping of the blades equivalent to a forwardinclination of the rotor cone and of the resultant force vector 8 inFig. l proportional to theangular speed. In

other words, the hinged lifting rotor produces a ampins'moment When theattitude of the aircratt is changed. proportional to the rate of changeof attitude. This is a desirable property f he otor.

It would be possible to avoid the instability with changes of attitudeof the aircraft if the o3 angle of the conventionalrotor would :beincreased to avalue somewhat below 9.0". In .such a rotor lql c flappingwould be almost entirely. suppressed. There-would he no instability withattitude. changes of the aircraft, but there would also be no stabilitywith speed changes and there would be no damping with rate of attitudechanges of the aircraft. The desirable properties would be eliminatedtogether with the undesirable properties.

'Ihe'rotor according to the invention, however, retains all thoseproperties of the conventional rotor with zero or small 63 angle wherecyclic flapping only is involved, because with the rotor according tothe invention cyclic flappin pro, duces only small changes in-bladepitch angles.

\ Since the stability with speed and the damping;

with rate of attitude changes of 'theaircraft both involve substantiallycyclic flapping only these two properties are for the rotor according tothe invention of thesameorder of magnitude as for the conventional rotorwithzero or small .63 angle. As to the desirable amount ,of the .53cycangle theoretical considerations indicate that probably best resultswill be obtained with a 63 an angle betweenlS and 30 depending: on thetype of blades. 7

.Ihe second of the above mentioned threerotor characteristics isfundamentally changed for the rotor according to the invention. Insteadof" being instable with respect to attitude changesof the aircraft, therotor according to theinvention is stable in the whole flight range. Thequalitative explanation for this stabili ty is as -fol-- lows.The-backward inclination. of therotor cone and of the resultant forcevector 8 increasesin forward night with increasing blade pitch angle 0or, in other words, a decrease in bladepitch .angle .0 produces aforward inclination of therotor cone. When changing the attitude of theaircraft by a nose-up motion, the lift on the; rotor-becomes larger,consequently the blades iii-- crease their collective flapping angle,which produces in the rotor accordingto -the inventionan appreciablereduction in blade pitch angle 0. lheforward inclination of the rotorcone and of the resultant force vector 8 accompanying this re?" ductionin blade pitch angle 0, overcorjnpensates the natural tendency of therotor cone to incline; backward because of the nose-up motion and thedisplacement of the resultant force vector 8-is a forward inclinationproducing a stabilizing mo-' ment.

A second very undesirable phenomenon of the conventional lifting rotorisits sensitivity to stall of the rotating wings, especially in pull-up maneuvers, when the resultant rotor force vector 8 in Fig. 1 is increasedabove the normal value. The rotor according to the invention responds toevery increase of rotor force vector ll, because it is ac? companied bya collective upward flapping of the blades, with a marked reduction inb-ladepitch angle '0 and, therefore, the stall limit is shifted" out ofthe normal operational range of the rotor.

The fulfilment of the third of the objects 9f" the invention, theprovision of windmil' ling in; power-off flight without adjustment ofblade pitch control, is obvious. As soon as the an gnlar speed of therotor-dropsbecause of reduction or cut-off of the driving power theblades fi'ap up wards in unison anda reduction in blade pitch angleoccurs which is sufiicient to keep. up the 179: h n t a i ghtiv e ueed uars eed- Another embodiment of the invention is shown' in-Figs."'7 and 8The inner links 31 are hinged tothe rotatingsha ft I so as to allowamotion of the" nner l nks 3! ab ut h r zonta axi fle 1. 1 li ks as ahin d to he i n links 2 as to allow horizontal motionsof the outer linksb ut t e ve tica axi h b ad l a etate' i connected to he Oiiter 1ink$ 3Qtya it ary g ivot so that they may etate about hei lo it nal. axi mianected to the l de T lade qi t ol er-fie while a second pair of linkcontrol hornsiij qnnecte to the o te i ks .33 One an o th alk eam 3 icon ted t t e n poin or. poi to a ta hme t o th l n was 1 35. Atitsotherend the wa1k=ing.beam 35, 1 y 5 avertical pushrod 31, theupper-end of this push: 1

rodbeing connected to the end point Ha of {the blade control horn ;l!.The vertical pushrods c4,

connected to the midpoint 38 of the walking-- beams 36. are, for thepurpose of rotor contnoi moved the vertical direction. As previous casesthe rotor pitch control will be disregarded n it ll be a sumed t att e pnts .3 are" w thou -v tica moti n. .113 o. cyclic flappin motionoppositeiblades'.

flap in opposite directions (see-saw motion) No change in blade pitchangle takes place for cyclic flapping because the walking beams 36 arefree to participate in the see-saw motion. The axes 32 are the cyclicflapping axes for the blades and in the rotor of Fig. 7 cyclic flappingproduces no change in blade pitch angle. By choosing for the push rods26 a point of attachment 38 different from the midpoint of the walkingbeam 36, any desired coupling between cyclic flapping angle and bladepitch angle may be obtained. The walking beam 35 will also be referredto as a tip path pFane follower because it tilts about its point ofattack 38 by an angle proportional to the tilting angle of the blade tippath plane thereby rendering the blade pitch substantiallynon-responsive to cyclic blade flapping.

When the blades flap in unison and the controls are held fixed, there isan appreciable decrease of blade pitch angle with increased flappingangle. If the point Ha were vertically fixed, an increase in flappingangle Aficoi of blade I would produce a decrease in blade pitch angle A6of Actually, however, the point Ila is not vertically fixed.

When the blade la flaps upwardly by an angle ABcol it moves the end 35aof the walking beam 35 in the upward direction. Since the walking beam35 is held in the midpoint 38 the other end Ha of the walking beam ismoved in the downward direction thereby reducing the pitch angle ofblade I by an amount equal to that indicated in the above equation. Thetotal decrease in blade pitch angle produced by collective flapping,therefore, is:

Acol= ABco1 tan 63 col The rotor of Fig. 7 is shown with hinges withvertical axes 34 because this embodiment of the invention lends itselfadvantageously to the addition of such hinges. The modern developmenttrend, however, is toward avoiding vertical hinges. In a two-bladedrotor the vertical hinges may be omitted if the rotor shaft is connectedto the frame of the aircraft with suflicient elasticity to alow forhorizontal motions of the hub, and it is assumed that in the cases shownin Fig. and in Figs. 9 to 11 such provisions are made. In three and morebladed rotors the vertical hinges may be omitted if the rotor hub is ofthe freely floating type and tiltably connected to the shaft as in Fig.6.

Another embodiment of the invention is shown in Figs. 9 and 10. Thelinks |2 are hinged to the rotor shaft I so as to rotate freely aboutthe axes |3. The blades 1 are rotatably connected to the links l2 bymeans of pitch varying pivots so as to rotate about their longitudinalaxes M. The blades I carry control horns I]. At the end points l8 of thecontrol horns H the vertical pushrods 40 are attached. Furthermore atthe points 39 of the control horns vertical push rods 4| are attached.The lower ends of the vertical pushrods 4| are connected by a crosshead42 which is supported at its midpoint 43 by a vertical link 44. Thevertical link 44 is for the purpose of collective blade pitch control,operated in a manner well known in the art and therefore not shown inthe drawing. For a fixed position of the collective blade pitch controlthe point 43 is vertically fixed.

The vertical push rods 40 are, for the purpose of cyclic pitch control,also operated in a manner well known in the art, except that norestraint must exist which prevents a unison vertical motion of bothpush rods 48. For neutral position of the controls the straight linethrough the points I8 is horizontal but a free vertical motion of thisline without angular displacement is possible.

For a cyclic flapping motion when opposite blades flap in oppositedirections (see-saw motion) the cross head 42 participates in the motionand the push rods 4| move freely up and down without restraining theblades. The points l8, however, in a cyclic flapping motion arevertically fixed.

For a collective flapping motion, when both blades flap in unison, thepushrods 40 move freely up and down without restraining the blades. The

points 39, however, in a collective flapping motion are verticallyfixed.

The axis 28 through the center point of the axis I3 and through thepoint l8 on the control horns ll is the cyclic flapping axis. The axis23 is, according to the invention, located so that a change in cyclicflapping angle produces a change in blade pitch angle which isappreciably smaller than the cyc ic flapping angle change.

The axis 30 through the center point of the axis l3 and through thepoint 39 on the control horn I! is the collective flapping axis. Theaxis 3|! is, according to the invention, located so that an increase incollective flapping angle produces a decrease in blade pitch angle whichis appreciably larger than the collective fiapping angle increase. Thepushrods 4B and 4| wi l also be referred to as vertical control links.The cross head 42 will be referred toas a tip path plane followerbecause it tilts about its mid point 43 by an angle proportional to thetilting angle of the b ade tip path plane thereby rendering the bladepitch substantially non-responsive to cyclic blade flapping.

Fig. 11 illustrates another embodiment of the invention. The walkingbeam 2| is hinged to the rotor shaft so as to allow free see-saw motionsabout the axis 22. The blades 1 are rotatably connected to the walkingbeam 2| by means of a pitch varying pivot extending in the direction:

of the axis 30 so as to al ow free motions about the axis 30. the axis22 and no changes in blade pitch are produced by cyclic flappingmotions. Collective flapping takes place about the axis 38, and anincrease in collective flapping angle produces a decrease in blade pitchangle which is appreciably larger than the collective flapping angleincrease.

While the location of the axis 22, as drawn in Fig. 11. corresponds to azero 63 m angle. a moderate value of 5: eye of 15 to 30 is possible anddesirable.

No means of rotor control are prov ded on the An aircraft with a liftingrotorv without rotor control must be maneuvered by rotor of Fig. 11.

separate means of control like auxiliary air screws which may not alwaysbe practical. The advantage of such an aircraft, however, lies in thevery simple construction of the lifting rotor. In

spite of this simplicity the rotor of Fig. 11 fulfills- :all the objectsof the invention.

I wish it to be understood that the constructions I have describedherein are shown by way of example and are not to be construed as theonly manners of carrying out the invention. It is. my intention to coverall modifications falling Y Cyclic flapping takes place about aoiaoer llwithin the inventive concept: as defined: by the appended claims.

I claim:

13.. an aircraft having an aircraft body, a lifting. rotor comprising a.center portion connected to said aircraft body, a hub like membertiltably connected tosaid, center portion, a plurality of blades, aflapping hinge for each of. said blades connected to said hub likemember, each of said blades being rotatably connected to its associatedflapping. hinge so as to allow rotation, of said blades about theirlongitudinal axes, a control: horn connected; in trailing relation toeach of said blades, and. an actuating element for effect ng positivepitch control attached to each control horn at a point of attachmentonsaid control horn, the line determined by each point of attachment andthe rotor center constituting a virtual cyclic flapping axis and theline determined by each point of attachment and the center of theassociated flapping hinge constituting a virtual collective flappingaxis, each of said virtual cyclic flapping axes forming an angle greaterthan 45 with the longitudinal'axis of its associated blade when saidblade-is radially disposed, and each of said virtual collective flappingaxes forming an angle less than 45 with the longitudinal axis of itsassociated blade when said longitudinal axis is radially disposed, saidangles being measured from the longitudinal axis of each blade in thedirection of rotation of the rotor.

2. In an aircraft having an aircraft body, a lifting rotor comprising a.center portion connected to said aircraft body, a plurality of blades, aflapping hingefor each of. said blades connected to said center portion,each of said blades being rotatably connected to its assoeiatedflap pinghinge so as to allow rotation of said blades about their longitudinalaxes, a control horn connected to each of saidblades, each havinga pointof attachment for cyclic pitch control only and defining a virtualcyclic flapping axis through said point of attachment and. throughtheoenter of its associated flapping hinge, each of saidv control hornshaving a'second point of attachment for collective pitch control onlyand defining avirtualcollective flapping axis through said second pointof attachment and through the center of its associated flapping hinge,an actuating element connected to each'of said horns at second point ofattachment, a tip path plane following element tiltable about a pivotintersecting the axis of said center portion and interconnecting saidactuating elements, each of said virtual cyclic flapping axes forming anangle greater than 45 degrees with the longitudinal axis of itsassociated blade when said blade is radially disposed, and each of saidvirtual collective flapping axes forming an angle less than 45 degreeswith the longitudinal axis of its associated blade when saidlongitudinal axis is radially disposed, said angles being measured fromthe longitudinal axis of each blade in the direction of rotation of therotor.

3. In an aircraft having an aircraft body, a

lifting rotor comprising a center portion rotatably connected to saidaircraft body, a plurality of blades, a hinge mechanism for each of saidblades effectively connecting said blades to said center portion topermit flapping motion of said blades and to permit rotation of saidblades-substantially about their longitudinal blade axis, ablade pitchcontrol horn connected to each of said blades having a point ofattachment for collec-' tive pitch control only and asecondzpoint ,of 75attachment for cyclic pitch control only, .collecg- 12 tive pitchcontrol links. connected? to thejfirs't point. of attachment of; saidpitch control hem and beingunrestrained in their cyclic motion; andcyclic pitch, control; links. connected to the second point ofattachment. of, said pitch control horn and being unrestrained .in theircollective motion, a tippath'plane following element tiltable about apivot intersecting the axis of said center portion and interconnectingsaid collective pitch, control links, said hinge mechanism, and said.-points of attachment, defining an effective substantially horizontalcollective flapping axisv and an effective substantially horizontalcyclic;;flapping axis for each of said blades, said collective flappingaxis passing through the center of. said hinge mechanism and throughsaid first pointof; attachment of said'pitch control horn and extendingoutwardly when seen in the direction of rotation of its associate blade,and said cyclicflapping axis passing through the center of said-; hingemechanism and through said second point; of attachment of said pitchcontrol horn and extending substantially in the direction of rota-- tionof the blade, each of said effective cyclic flapping axes formingsubstantially a right angle with the longitudinal axis of its associatedblade, and each of said effective collective flapping-axes forming anacute angle with the longitudinal axis of its associated blade when saidlongitudinal axis is radially disposed, said angles being measured fromthe longitudinal axis of each blade in the direction of rotation of therotor, whereby the blade pitch is made substantially responsive tocollective blade flapping only and substantially non-responsive tocyclic blade flapping.

4. In an aircraft having'an aircraft body, a lifting rotor comprising acenter portion connected to said aircraft body, a hub like membertiltablyconnected to said center portion, a plurality of outer flappinghinges connected to said hub like member, a blade rotatably-connectedto. each of said outer flapping hinges so as to allow: rotation of saidblades about their longitudinal axes, control horn connected to each ofsaidblades,

and an actuating element for effecting positive. pitch control attachedto each: control. horn. at

a point of attachment on said control born, the line determined by eachpoint of attachment and the rotor center constituting a virtual cyclic.flap ping axis and the line determined by each point of attachment andthe center of the associated flapping hinge. constituting a virtualcollective flapping axis, each of said virtual cyclic flapping axesforming an angle greater than 45 with the longitudinal axis of itsassociated blade whensaid bladeis radially disposed, and each of-saidvirtual collective flapping axes forming an angle less than 45 with thelongitudinal axis of its associated blade when saidv lonigtudinal axisis; radially disposed, said angles. being measured. from thelongitudinal axis of each blade in the direction of rotation of therotor.

5. In an aircraft having an aircraft body, a lift-i ing rotor comprisinga center portion connected to'said aircraft body, a hub likemember'tiltably connected to said center portion, a plurality of outerfiappinghinges connected to said .hub like member, a blade rotatablyconnected to each of said outer flapping hinges $085 to allow rotationofsaid blades about their longitudinal axes. a control horn connected intrailing relation to each of said blades, and-an actuating element foreffecting positive pitch control attachedto each control horn at a pointof attachment on said; control born, the line determined by each-pointof attachment and the rotor center constituting a virtual cyclicflapping axis and the line determined by each point of attachment andthe center of the associated flapping hinge constituting a virtualcollective flapping axis, each of said virtual cyclic flapping axesforming substantially a right angle with the longitudinal axis of itsassociated blade when said blade is radially disposed, and each of saidvirtual collective flapping axes forming an acute angle with thelongitudinal axis of its associated blade when said longitudinal axis isradially disposed, said angles being measured from the longitudinal axisof each blade in the direction of rotation of the rotor, whereby theblade pitch is made substantially responsive to collective bladeflapping only and substantially non-responsive to cyclic blade flapping.

6. In an aircraft having an aircraft body, a lifting rotor comprising acenter portion connected to said aircraft body, a plurality of blades,flapping and pitch varying mechanism effectively connecting said bladesto said center portion, a flapping hinge for each of said blades andincluded in said mechanism, each of said blades being rotatablyconnected to its associated flapping hinge so as to allow rotation ofsaid blades about their longitudinal axes, a control horn included insaid mechanism and connected to each of said blades, actuating links foreffecting pitch control included in said mechanism and attached to saidcontrol horns, a tip path plane following element included in saidmechanism and interconnecting said blades with each other and tiltableabout a pivot intersecting the axis of said center portion, said elementand said blades performing together cyclic flapping motionssubstantially without changing the pitch of said blades, said mechanismrendering the blade pitch responsive to collective blade flapping withrespect to the plane of said element, whereby an increase in collectiveblade flapping angle produces a decrease in collective blade pitch anglewhich is larger than the increase of said collective blade flappingangle.

7. In an aircraft having an aircraft body, a lifting rotor comprising acenter portion connected to said aircraft body, a hub like member, acentral hinge mechanism connecting said hub like member with said centerportion to permit tilting motions of said hub like member, a pluralityof blades, outer flapping and pitch varying mechanism for each bladeeffectively connecting it to said hub like member, said outer flappingand pitch varying mechanism including an outer hinge for each of saidblades permitting each blade to flap with respect to said hub likemember, a pitch varying pivot for each of said blades permitting eachblade to rotate substantially about its longitudinal axis, a pitchcontrol horn connected to each blade having a point of attachment forpitch control, the line determined by each point of attachment and thecenter of said central hinge mechanism constituting a virtual cyclicflapping axis and the line determined by each point of attachment andthe center of said outer hinge constituting a virtual collectiveflapping axis, each of said virtual cyclic flapping axes formingsubstantially a right angle with the longitudinal axis of its associatedblade, and each of said virtual collective flapping axes forming anacute angle with the longitudinal axis of its associated blade when saidlongitudinal axis is radially disposed, said angles being measured fromthe longitudinal axis of each blade in the direction of rotation of therotor, said hub like member and said blades performing together cyclicflapping motions substantially without changing the pitch of saidblades, said outer flapping and pitch varying mechanism rendering theblade pitch responsive to collective blade flapping with respect to theplane of said hub like member.

KURT H. HOHENEMSER.

References Cited in the file of this patent UNITED STATES PATENTS NumberName Date 2,045,355 Hays June 23, 1936 2,086,802 Hays July 13, 19372,192,492 Bennett Mar. 5, 1940 2,397,154 Platt Mar. 26, 1946 2,429,646Pullin Oct. 28, 1947 FOREIGN PATENTS Number Country Date 476,596 GreatBritain Dec. 13, 1937 OTHER REFERENCES Ser. No. 254,867, Flettner (A. P.C.) published May 25, 1943.

